1. Field of the Invention
This invention relates to a turbine rotor blade for a gas turbine engine, to a turbine rotor incorporating such blades, and to a gas turbine engine comprising such a rotor.
2. Brief Description of the Related Art
The turbine of a gas turbine engine depends for its operation on the transfer of energy between the combustion gases and turbine. The losses which prevent the turbine being totally efficient are due in part to leakage flow of working fluid over the turbine blade rotor tips.
In turbines with unshrouded rotor blades, a portion of the working fluid flowing through the turbine tends to migrate from the pressure surface to the suction surface of the blade aerofoil through the gap between its tip and the stationary shroud or casing. This leakage occurs because of a pressure difference that exists between the pressure and suction sides of the aerofoil. The leakage flow also causes flow disturbances to be set up over a large portion of the span of the aerofoil, which also leads to losses in the efficiency of the turbine.
By controlling the leakage flow of air or gas across the tips of the blades it is possible to increase the efficiency of each rotor stage. One solution is to apply a shroud to the rotor tip. When the rotor blades are assembled together in the disc that carries them, these shrouds form a continuous ring that prevents the leakage flow from the aerofoil pressure to suction side at the tip. There is still an axial leakage through the gap between the casing and the rotating shroud, but the penalties in terms of aerodynamic losses are much reduced—often helped by the inclusion of a form of labyrinth seal on the shroud top.
However, the rotating shroud has a large weight penalty. As a result, the aerofoil blade speed may be constrained, to achieve acceptable blade stresses. This, though, will have the effect of increasing the aerodynamic loading that also results in reduced efficiency, negating some of the benefit of the shroud.
The use of a shroud ring is made more difficult if the turbine blades also operate at very high temperatures, desirable in helping to achieve high thermal efficiencies. These temperatures are limited by the turbine vane and blade materials. Cooling of these components is necessary to achieve acceptable component life, which is a function of the material temperature, stresses and material properties.
A large number of cooling systems are now applied to modern gas turbine blades. Such cooling systems are described for instance in Cohen H, Rogers G F C, Saravanamuttoo H I H, 1981, “Gas Turbine Theory”, p. 232–235, Longman, and Rolls-Royce plc, 1986, “The Jet Engine”, p. 86–88, Renault Printing Co Ltd. The more cooling that is provided, the lower the resulting material temperatures, and thus the higher the blade stresses allowable for a given component life. Cooling is achieved using relatively cool fluid bled from the upstream compressor system, bypassing the combustion chamber between the last compressor and first turbine. This air is introduced into the turbine blades where cooling is effected by a combination of internal convective cooling and external cooling. However, this cooling comes at a penalty. Its use penalises the overall efficiency of the machine, and as a result the turbine designer tries to minimise the quantity of cooling air used. All of these design constraints often leads to rotor blades of first staged turbines being shroudless—the extra weight and thus higher stresses caused by a shroud ring simply cannot be accommodated. However, ways of reducing the high aerodynamic penalty of the resulting tip leakage flow continue to be sought.
U.S. Pat. No. 5,525,038 discloses a blade aerofoil design intended to reduce tip leakage losses. In that document, the tip region of the suction side of the aerofoil has a bowed surface with an arcuate shape. The arcuate shape of the bowed surface has progressively increasing curvature toward the tip of the rotor blade, so that a radial component of a normal to the suction side bowed surface becomes progressively larger toward the tip. It is to be noted that the aerofoil has a bowed surface in the tip region extending chordally all the way from the leading edge to the trailing edge. In addition, all the leans of the tip described, whether tangential and/or axial, are applied to the whole of the tip region.
A particular area of the blade that requires attention in its design is the trailing edge. Preferably, this is kept thin to minimise aerodynamic losses, but as a result it is difficult to cool, and tensile stresses have to be minimised. Cooling is achieved by films of air ejected upstream of the trailing edge on to the aerofoil surfaces, and by drilling cooling holes into the trailing edge fed from larger radial passages within the main body of the aerofoil. The aerofoil disclosed in U.S. Pat. No. 5,525,038 thus has some disadvantages: firstly, the curved trailing edge cannot easily have cooling holes machined into it. Ideally, this is done in one operation to minimise cost, but this requires all holes to lie in the same plane, i.e. the trailing edges of the aerofoil sections making up the blade have all to lie in one plane. A curved trailing edge (with a progressively increasing curvature) will require holes to be machined in multiple operations, incurring significant extra cost. Secondly, the leant tip will give rise to additional bending stresses in the blade. In the main body of the blades these can usually be accommodated by changes in the detailed design, such as increasing wall thickness locally. However, this cannot be done in the trailing edge region, there will simply be higher stresses in it. This will result either in a reduced life of the component, or require additional cooling, which will impair the performance of the engine.